1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically for an air cooled turbine blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with one or more stages or stator vanes and rotor blades that react with a hot gas stream and produce mechanical work. The first stage airfoils (vanes and blades) are exposed to the highest temperature gas flow and therefore require the most cooling. In order to allow for higher turbine inlet temperatures—and therefore higher engine efficiencies—better cooling is required if material properties are not advanced enough. Also, since the airfoil cooling air is typically bled off from the compressor, the cooling air used does not contribute to producing any work in the engine. It is a design objective to not only provide for better cooling capability, but also to use a minimal amount of cooling air to higher efficiency.
FIG. 1 shows a first stage blade external pressure profile. A forward region of the leading edge and the pressure side surface experiences a high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than on the pressure wall side. Thus, a near wall serpentine flow blade cooling design can be divided into four zones: 1) the blade leading edge region, 2) the blade pressure side section, 3) the blade suction side section, and 4) the blade trailing edge region. Each individual cooling zone can be independently designed based on the local aerodynamic pressure loading conditions. dividing the airfoil into these four zones increases a design flexibility to redistribute cooling flow and/or add cooling flow for each zone and therefore increase a growth potential (use the similar design for larger airfoils) for the cooling design. Also, individual serpentine flow circuits used in each zone can further enhance the flexibility of the cooling flow distribution. With this design approach, a more uniform temperature distribution for the airfoil mid-chord section can be achieved. A uniform temperature distribution will reduce hot spots from appearing on the airfoil that causes erosion and short blade life.
FIG. 2 shows a first stage blade external heat transfer coefficient profile. The airfoil leading edge, the suction side immediately downstream from the leading edge, and the airfoil trailing edge region experiences the higher hot gas side external heat transfer coefficient than does the mid-chord section of the pressure side and downstream of the suction side. The heat load for the airfoil aft section is higher than the forward section. This heat load distribution can also be subdivided into four zones as in the above described pressure profile of FIG. 1. Individual zones can then be designed based on the local heat load to achieve a uniform metal temperature distribution profile. Different cooling channel size for each zone can be used to adjust for the required cooling flow rate to achieve the metal temperature level.